XFOIL         Version 6.94
  
 Calculated polar for: GOE 174 (ALBATROS 5020) AIRFOIL                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.5926   0.01158   0.00352  -0.1126  0.6223  0.0390
   0.500   0.6471   0.01147   0.00329  -0.1122  0.6008  0.0438
   1.000   0.7012   0.01136   0.00316  -0.1117  0.5778  0.0634
   1.500   0.7546   0.01064   0.00331  -0.1116  0.5526  0.4037
   2.500   0.8581   0.01000   0.00338  -0.1095  0.4890  1.0000
   3.000   0.9100   0.01042   0.00356  -0.1088  0.4533  1.0000
   3.500   0.9610   0.01097   0.00385  -0.1080  0.4230  1.0000
   4.000   1.0116   0.01151   0.00418  -0.1071  0.3887  1.0000
   4.500   1.0578   0.01233   0.00463  -0.1057  0.3220  1.0000
   5.000   1.1058   0.01312   0.00514  -0.1045  0.2738  1.0000
   5.500   1.1505   0.01427   0.00583  -0.1030  0.1974  1.0000
   6.000   1.1818   0.01688   0.00760  -0.0997  0.0577  1.0000
   6.500   1.2210   0.01854   0.00907  -0.0972  0.0063  1.0000
   7.000   1.2657   0.01950   0.01017  -0.0954  0.0061  1.0000
   7.500   1.3077   0.02063   0.01152  -0.0932  0.0064  1.0000
   8.000   1.3476   0.02184   0.01294  -0.0907  0.0071  1.0000
   8.500   1.3818   0.02340   0.01476  -0.0873  0.0079  1.0000
   9.000   1.4083   0.02515   0.01677  -0.0828  0.0088  1.0000
   9.500   1.4202   0.02773   0.01964  -0.0767  0.0099  1.0000
  10.000   1.4269   0.03097   0.02315  -0.0712  0.0109  1.0000
  10.500   1.4198   0.03604   0.02850  -0.0666  0.0121  1.0000
  11.000   1.4137   0.04188   0.03459  -0.0639  0.0131  1.0000
  11.500   1.4028   0.04870   0.04163  -0.0613  0.0147  1.0000
  12.000   1.4241   0.05188   0.04481  -0.0545  0.0183  1.0000